Cmc articles having small complex features

ABSTRACT

A ceramic matrix composite (CMC) component for gas turbine engines, the component having fine features such as thin edges with thicknesses of less than about 0.030 inches and small radii of less that about 0.030 inches formed using the combination of prepreg plies layed up with non-ply ceramic inserts. The CMC components of the present invention replace small ply inserts cut to size to fit into areas of contour change or thickness change, and replace the small ply inserts with a fabricated single piece discontinuously reinforced composite insert, resulting in fewer defects, such as wrinkles, and better dimensional control.

FIELD OF THE INVENTION

The present invention relates generally to ceramic matrix turbine enginecomponents, and more particularly, to a ceramic matrix composite gasturbine engine component having small complex features.

BACKGROUND OF THE INVENTION

In order to increase the efficiency and the performance of gas turbineengines so as to provide increased thrust-to-weight ratios, loweremissions and improved specific fuel consumption, engine turbines aretasked to operate at higher temperatures. The higher temperatures reachand surpass the limits of the material comprising the components in thehot section of the engine and in particular the turbine section of theengine. Since existing materials cannot withstand the higher operatingtemperatures, new materials for use in high temperature environmentsneed to be developed.

As the engine operating temperatures have increased, new methods ofcooling the high temperature alloys comprising the combustors and theturbine airfoils have been developed. For example, ceramic thermalbarrier coatings (TBCs) have been applied to the surfaces of componentsin the stream of the hot effluent gases of combustion to reduce the heattransfer rate, provide thermal protection to the underlying metal andallow the component to withstand higher temperatures. These improvementshelp to reduce the peak temperatures and thermal gradients of thecomponents. Cooling holes have been also introduced to provide filmcooling to improve thermal capability or protection. Simultaneously,ceramic matrix composites have been developed as substitutes for thehigh temperature alloys. The ceramic matrix composites (CMCs) in manycases provide an improved temperature and density advantage over metals,making them the material of choice when higher operating temperaturesand/or reduced weight are desired.

A number of techniques have been used in the past to manufacture hotsection turbine engine components, such as turbine airfoils usingceramic matrix composites. One method of manufacturing CMC components,set forth in U.S. Pat. Nos. 5,015,540; 5,330,854; and 5,336,350;incorporated herein by reference in their entirety and assigned to theassignee of the present invention, relates to the production of siliconcarbide matrix composites containing fibrous material that isinfiltrated with molten silicon, herein referred to as the Silcompprocess. The fibers generally have diameters of about 140 micrometers orgreater, which prevents intricate, complex shapes having features on theorder of about 0.030 inches, such as turbine blade components for smallgas turbine engines, to be manufactured by the Silcomp process.

Other techniques, such as the prepreg melt infiltration process havealso been used. However, the smallest cured thicknesses with sufficientstructural integrity for such components have been in the range of about0.030 inch to about 0.036 inch, since they are manufactured withstandard prepreg plies, which normally have an uncured thickness in therange of about 0.009 inch to about 0.011 inch. With standard matrixcomposition percentages in the final manufactured component, the use ofsuch uncured thicknesses results in final cured thicknesses in the rangeof about 0.030 inch to about 0.036 inch for multilayer ply components,which is too thick for use in small turbine engines.

Complex CMC parts for turbine engine applications have been manufacturedby laying up a plurality of plies. In areas in which there is a changein contour or change in thickness of the part, plies of different andsmaller shapes are custom cut to fit in the area of the contour changeor thickness change. These parts are laid up according to a complicated,carefully preplanned lay-up scheme to form a cured part. Not only is thedesign complex, the lay-up operations are also time-consuming andcomplex. Additionally, the areas of contour change and thickness changehave to be carefully engineered based on ply orientation and resultingproperties, since the mechanical properties in these areas will not beisotropic. Because the transitions between plies along contourboundaries are not smooth, these contours can be areas in whichmechanical properties are not smoothly transitioned, which must beconsidered when designing the part and modeling the lay-up operations.

FIG. 1 depicts an exemplary uncoated airfoil (uncooled) 10. In thisillustration the airfoil 10 comprises a ceramic matrix compositematerial. The airfoil 10 includes an airfoil portion 12 against which aflow of gas is directed. The airfoil 10 is mounted to a disk (not shown)by a dovetail 14 that extends downwardly from the airfoil portion 12 andengages a slot of complimentary geometry on the disk. The airfoil 10does not include an integral platform. A separate platform can beprovided to minimize the exposure of the dovetail 14 to the surroundingenvironment if desired. The airfoil has a leading edge section 18 and atrailing edge section 16. Such a composite airfoil is fabricated bylaying up a plurality of plies.

FIG. 2 is a prior art illustration (perspective) of how such a compositeairfoil of FIG. 1 has been laid up. FIG. 3 represents a front view ofthe lay-up of these pre-preg plies. The airfoil 10 comprises a pluralityof pre-preg plies 40 arranged around a centerplane 24. There are anumber of root (pre-preg) plies 41 and smaller (pre-preg) plies 42arranged between larger (prepreg) plies 40, 44. Referring back to FIG.1, the smaller plies, in particular root plies 41, are required toprovide the dovetail geometry. In addition, each of the plies 40includes tow that is oriented in a predetermined direction. Of course,care must be taken to not only provide the proper size ply in the properlocation, but also to properly orient the tow direction of each of theplies.

Still other techniques attempt to reduce the thickness of the prepregplies used to make up the multi-layer plies by reducing the thickness ofthe fiber tows. Theoretically, such processes could be successful inreducing the ply thickness. However, practically, such thin plies aredifficult to handle during part manufacturing, even with automatedequipment. This can result in wrinkling of the thin plies, amanufacturing defect that can result in voids in the article, and adeterioration of the mechanical properties of the article, and possibleply separation. In addition, problems arise, as airfoil hardwarerequires the ability to form small radii and relatively thin edges,features required in smaller articles, such as narrow chord airfoils.The high stiffness of the fibers, typically silicon carbide, in theprepreg tapes or plies, can lead to separation when attempting to formthe plies around tight bends and corners with small radii. The fibercoatings may also crack or be damaged. This leads to a degradation inthe mechanical properties of the article in these areas with resultingdeterioration in durability.

What is needed is a method of manufacturing CMC turbine enginecomponents that permits the manufacture of features having a thickness,particularly at the edges, in the range of about 0.015 inch to about0.021 inch, as well as small radii, the radii also in the range of lessthan about 0.030 inches. In addition, a method of manufacturing CMCturbine engine components having features with a thickness less thanabout 0.021 inch is also needed.

SUMMARY OF THE INVENTION

The present invention is a ceramic matrix composite (CMC) component foruse at high temperatures, such as in gas turbine engines, the componentshaving fine features such as thin edges with thicknesses of less thanabout 0.030 inches and small radii of less that about 0.030 inchesformed using the combination of prepreg plies layed up with non-plyceramic inserts. Turbine components produced using the processes of thepresent invention minimize the use of ply inserts cut to size to fitinto areas of contour change or thickness change, particularly inapplications in which there is a significant thickness change over ashort distance, wherein, the short or small distances are measured in adirection substantially transverse to the direction of changes inthickness. These thickness changes are considered substantial in plylay-ups when the change in thickness is as little as 10%, since failureto properly design a component to account for such changes can result indefects in the finished component.

Current practices require cutting small ply inserts to size to enablenet shape lay-up, which is necessary to minimize compaction during cureresulting in fewer defects, such as wrinkles, and better dimensionalcontrol.

Turbine components are modeled using non-ply ceramic inserts incombination with prepreg layers in the present invention. The compositecomponent comprises a lay-up of substantially continuous plies, each plyin the lay-up of substantially continuous plies having a plurality oftows extending substantially parallel to each other in an uncured matrixmaterial, each ply being positioned so that the tows extend at apreselected angle to the tows in an adjacent ply. Non-ply ceramicinserts, as used herein, means both discontinuously reinforced compositeinserts and monolithic ceramic inserts. The components are modeled usingprepreg plies or tapes in combination with the non-ply ceramic inserts.In areas where complex features are present, non-ply ceramic inserts areincorporated into the component, so that the turbine component is acombination of prepreg layers and non-ply ceramic inserts. Althoughprepreg plies may be cut to a smaller size and included in combinationwith substantially full length prepreg layers and the non-ply ceramicinserts, the non-ply ceramic inserts are modeled into the component toreplace a substantial number of the small prepreg plies that previouslywere cut to size to provide for a change in thickness or a change incontour, the replacement of which provides a predetermined shape. Theinsert or discontinuously-reinforced composite portion is adjacent tothe reinforced ceramic matrix composite portion comprising the pluralityof continuous tows It is cured to the reinforced ceramic matrixcomposite portion to form the component.

The non-ply ceramic insert or piece is designed and produced to minimizethe number of small fiber plies, cut and inserted into a portion of acomponent to allow for a change in thickness or contour, whilemaximizing the number of continuous fiber plies that extend along thesubstantially full length of the component. A non-ply ceramic insertsmay include a plurality of configurations. The discontinuouslyreinforced composite insert may be made by cutting pre-preg plies intosmall pieces, mixing the small pieces with a slurry of matrix materialto form a paste or putty. Lengths of fiber or tow may be substituted forthe cut plies or may be used along with and in addition to the cutplies. The paste or putty is applied into cavities of the component, asit is layed up, forming an uncured insert, which cures on drying.Alternatively, the mixture can be molded and cured to form a curedinsert, which is assembled into the component. Inserts made fromdiscontinuously reinforced composite, while having properties that arenot quite isotropic, nevertheless are less directional than a cured CMClay-up. These mechanical properties are referred to herein as“substantially isotropic,” since they are not quite isotropic, but arenot directional.

An advantage of the present invention is that a plurality of small, cutfabric plies can be replaced by a single discontinuously reinforcedcomposite insert. The discontinuously reinforced composite insert can beprovided as a material having substantially isotropic properties.

Another advantage of the present invention is that manufacture of anaircraft engine component can be simplified by elimination of a complex,time-consuming lay-up scheme.

Yet another advantage of the present invention is that the use ofdiscontinuously reinforced composite inserts will allow for theinclusion of fine features, such as thin sections and small radii,without compromising the mechanical properties of the component.

Other features and advantages of the present invention will be apparentfrom the following more detailed description of the preferredembodiment, taken in conjunction with the accompanying drawings whichillustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 depicts a CMC airfoil for use in a gas turbine engine.

FIG. 2 depicts a prior art method for laying up the CMC airfoil of FIG.1.

FIG. 3 depicts a front view of the lay-up of FIG. 2.

FIG. 4 depicts an insert for use in the present invention replacing theplies set forth in FIG. 3.

FIG. 5 depicts an airfoil of the present invention layed up with insertsand pre-preg plies.

FIG. 6 depicts a lay-up of a narrow cord blade, in cross-section, theblade having a trailing edge insert and a rib inserts.

FIGS. 7A. 7B and 7C depict inserts of the present invention positionedat contour changes and at thickness changes over short distances.

DETAILED DESCRIPTION OF THE INVENTION

The present invention provides an aircraft engine component made of aCMC. The component comprises a plurality of substantially continuouspre-preg plies that extend substantially the length of the component. Atleast one discontinuously reinforced composite insert is incorporatedinto the component, the discontinuously reinforced composite inserthaving substantially isotropic properties. The discontinuouslyreinforced composite insert may extend substantially the length of thecomponent, but may be modeled to replace specially cut, smaller pre-pregplies at contours and at changes in discontinuously reinforced compositepart thickness.

As used herein, a fiber means the smallest unit of fibrous material,having a high aspect ratio, having a diameter that is very smallcompared to its length. Fiber is used interchangeably with filament. Asused herein, a tow means a bundle of continuous filaments. As usedherein, matrix is an essentially homogenous material into which othermaterials, fibers or tows specifically, are embedded. As used herein, apre-preg-ply, or simply pre-preg, means a sheet of unidirectional tow,or short lengths of discontinuous fiber impregnated with matrixmaterial, the matrix material being in resin form, partially dried,completely dried or partially cured. As used herein, a preform is alay-up of pre-preg plies that may or may not include an additionalinsert, into a predetermined shape prior to curing of the pre-pregplies.

The present invention is depicted as an insert 110 in FIG. 4. Thedepicted insert is a discontinuously reinforced composite materialhaving substantially isotropic properties in its preferred embodiment.Substantially isotropic properties may deviate slightly from beingexactly identical in all directions, but are distinguishable fromanisotropic properties, which are clearly different, that is, havingmechanical properties that are distinctly different directions at apoint in a body of it. Stated differently, an anisotropic material hasno planes of material symmetry. The discontinuously reinforced compositeinsert 110 can be manufactured by any convenient method. Thediscontinuously reinforced composite insert is not comprised of laid-upplies of material, but rather is a block of material that may have apredetermined shape, and that can be handled as an individual piece.Insert 110 may be fully cured or partially cured and then machined tothe predetermined shape. The discontinuously reinforced composite insertmay be fully dense or partially dense. If partially dense, as willbecome evident, the insert can be made fully dense as part of theoperations in forming a turbine engine component.

The insert may be formed by mixing chopped fiber with a matrix material.A variant utilizes chopped tow, chopped pre-preg plies, or chopped pliesthat are cured or partially cured. Typically, a coating selected fromthe group consisting of boron nitride, silicon nitride, silicon carbideand combinations thereof as is known in the art is applied to the fiber.This material is thoroughly mixed with matrix material to form a slurry,which can have a discontinuously reinforced composite viscosity rangingfrom a fluid to a thick paste. The material can be molded by anyconvenient means into a final shape or intermediate shape and cured. Thecured part can be final machined into a predetermined shape ifnecessary. Examples of shape-forming techniques include extrusion,casting, injection molding and pressing methods. If used as a paste orslurry, the material that forms the insert may be applied to areas ofthe preform that lacks material. In this circumstance, it may benecessary for the preform to provide support for the uncured paste orslurry. If this cannot be done, the formulation can be adjusted, as isknown, with polymer additions or sub-micron powder, to form athixotropic composition that is self-supporting. Chopped tow or filamentlengths used for either paste or slurry typically ranges from about 0.1inch to 1 inch. The fiber loading typically ranges from about 10% byvolume to about 50% by volume. These parameters are determined by themechanical properties requirements of the article, as well as to allowmolten silicon densification throughout the component.

The discontinuously reinforced composite insert is used in conjunctionwith a lay-up of plies for forming a turbine engine component. Theinsert is assembled with the plies and maintained with the plies as thecomponent is cured. If a fully integrated bond is desired, a number ofoptions are available, the option to be selected depending upon ease ofobtaining the desired bond. Thus, the insert itself may be a partiallycured molded article that can bond with the plies in the lay-up for thecomponent, the partially cured preform bonding with the resin matrix ofthe pre-preg plies during curing of the component. The insert may becarbon rich to facilitate a diffusion bond integral with the CMC matrixportion, the integral bond formed during molten silicon infiltration.Alternatively, when applied as a paste, the insert material can curewith the resin matrix of the pre-preg plies during curing of thecomponent. The final result is a fully dense turbine engine componenthaving at least two distinct portions, a cured reinforced ceramic matrixcomposite portion comprising a plurality of continuous tows extendingsubstantially parallel to each other in a matrix; and adiscontinuously-reinforced composite portion having substantiallyisotropic properties located at regions of contour changes and thicknesschanges of the component. The discontinuously reinforced compositeportion comprises discontinuous fiber-including material in a matrixmaterial. The discontinuously reinforced composite portion is adjacentto the reinforced ceramic matrix composite portion and is cured to thereinforced ceramic matrix composite portion. However, the use of theinsert permits the formation of very tight radii, and/or to form thinsections that were not achievable with prior art plies laid up to formthe thin section of, for example, a narrow chord airfoil. Furthermore,the formation of discontinuously reinforced composite inserts or the useof the insert material as a paste eliminates the prospect of wrinkling,and related defects as a result of handling a large number of small,thin plies.

The present invention is depicted in FIG. 5 as an alternate method ofmanufacturing the airfoil of FIG. 1. In one embodiment, this inventionenvisions replacing root plies 41 and smaller plies 42 with adiscontinuously reinforced composite insert 110 of FIG. 4 shown asinserts 510, 520, 530, 540, 550, 560, and 570 in FIG. 5. The insert 110preferably has substantially isotropic properties. The inserts depictedin FIG. 5 as 510, 520, 530, 540, 550, 560, and 570 replace the plies inFIG. 3 located at B, C, D, E, F, G and H respectively. The remainingplies in the preform are inserted prior to curing. The plies extend thefull length or substantially the full length of the component, theorientation of each of the plies being determined to provide therequired mechanical properties for the component, here an airfoil. Thus,a 0° orientation refers to a ply that is laid up so that the line offiber tows is substantially parallel to a preselected plane of thecomponent, for example the long dimension or axis of a turbine blade. A90° orientation refers to a pre-preg ply oriented at substantially 90°to the preselected plane. The remaining plies may be laid up in analtering formation, such as ±45° to the preselected plane of the part.Thus, for example, a sequence of plies is laid up in a sequence of 0°,+45°, −45°, 90°, 45°, +45° so that the component has tensile strength indirections other than along the axis. In this manner, the strength ofthe component can be modified to be directional (anisotropic) orsomewhat isotropic as desired. For the article depicted in FIG. 5, thefinal cured component is a CMC having tows extending in preselectedorientations, the plies which extended the full length of the componentor substantially the full length of the component yielding tows in amatrix extending substantially parallel to each other as a group. InCMCs having a plurality of plies, the cured component yields a pluralityof groups of continuous tows, the tows in each group extendingsubstantially parallel to each other in a matrix, each group oriented ata preselected angle to the tows in at least one other group and eachgroup having substantially anisotropic properties. However, the insertsadjacent to plies have substantially isotropic properties.

In an alternate embodiment of the present invention, the inserts areused to provide significant thickness changes over a short distance in athinner cross-section airfoil than is currently available using existingplies. FIG. 6 depicts two applications of the present invention for usein a narrow chord turbine blade 610 in which a portion of the thintrailing edge 612 includes a discontinuously reinforced composite insert650 having substantially isotropic properties. The blade also includespremolded rib inserts 680, made of the discontinuously reinforcedcomposite of the present invention.

The narrow chord turbine blade of FIG. 6 depicts a pair of airpassageways 614 that are fabricated into blade 610. Both the trailingedge insert 650 and rib inserts 680 are prefabricated using thediscontinuous materials set forth in the present invention. Both insert650 and inserts 680 replace a plurality of small plies that areextremely difficult to handle during lay-up operations. Inserts 650, 680are molded to near-net shape and machined to final dimensions afterbeing compacted and cured to remove the volatiles.

As shown in FIG. 6, insert 650 is positioned within the trailing edgereplacing a plurality of small plies that would be required to fill thegap between full length plies having a first end 652 on the suction side654 a second end 656 on the pressure side 658 of the blade. As shown inFIG. 6, insert 650 is bounded by three full-length plies extending fromthe suction side 654 to the pressure side 658. As used herein fulllength means that the plies extend the height of the blade from top tobottom, FIG. 6 being a cross-section through the height, extending intoand out of the plane of FIG. 6. It is envisioned that insert 650 can bemade somewhat larger (increased in cross-section) than shown in FIG. 6,thereby allowing replacement of at least one of the depicted full lengthplies. At least one full-length ply having a first end 652 and a secondend 656 is required on the suction side 654 and the pressure side 658.The direction of maximum stress in each blade design is known, and atleast one ply is oriented, typically on the outside of the insert, sothat its fibers run in a direction parallel to the direction of maximumstress. Each ply is of standard thickness of about 10 mils, comprising aplurality of unidirectional tows. However, if additional strength isneeded in directions offset from the direction of maximum stress, theinserts permit the substitution of thinner plies. These plies use thin,unidirectional tows, allowing ply thicknesses of less than 10 mils,generally from 5 mils to 9 mils. Although these thin plies are difficultto handle, they can be accommodated by the manufacturing process becausethey are full length plies that are laid up against a full length insertand used in limited numbers, replacing only one or two plies of standardthickness. Of course, insert 650 is increased in size proportionally toaccount for the difference in ply thickness when such thin plies aresubstituted for plies of standard thickness.

Inserts 680 are provided solely to replace the plurality of small pliesused at the change in thickness between air passages 614. As should beobvious, the lay-up of plies in this area requires many small plieshaving different widths that must be precisely placed. The fabricationof inserts 680 using the materials and methods of the present inventionand placement of the insert during lay-up is substantially easier andless prone to manufacturing error requiring scrapping of a cured bladethan laying up of a plurality of small plies.

To manufacture a blade such as the blade depicted in FIG. 6, continuousplies 652 along the suction side 654 are laid up on a lay-up tool. Amandrel (not shown) having inner wrap plies 690 is then placed at theappropriate locations, here where the air passageways are formed, asshown in FIG. 6. The outer plies extending from first end 652 on thesuction side 654 to second end 656 on the pressure side 658 are wrappedover to complete the lay-up sequence. The laid up blade is then curedunder pressure at temperature to remove the volatiles and to fullyconsolidate the blade. After consolidation and curing, the mandrel isremoved to provide air passageways 614. The matrix plies are placed overthe outer surface of the insert pieces 650, 680 to enhance bondingbetween the insert pieces and the continuous plies. Depending on thetechnique used, as discussed above, the blade can then be densifiedusing the melt infiltration process.

EXAMPLE 1

A slurry was prepared by utilization of SiC/SiC unidirectional prepregtape that is coated silicon carbide tow in a silicon carbide matrix. Thefibers comprising the tow typically are coated with a debond coatingsuch as boron nitride. The backing was removed from the prepreg byexposing the fabric to liquid nitrogen. The fabric was then cut intopieces having a size of about ¼ square in. and smaller. A proprietarysolution of Cotronics Resbond 931, a high temperature ceramic graphiteadhesive resin available from Cotronics Corp. of Brooklyn N.Y. andacetone was prepared by mixing with an equal weight of acetone. Thechopped pre-preg, about 3 g, was added to the solution in a weight ratioof about 3:1 pre-preg: solution to form a mixture. The mixture wasblended by a suitable means to achieve a uniform consistency. This canbe achieved by shaking, stirring, ball milling or other mixingtechniques. The viscosity of the mixture can be adjusted as requiredconsistent with its intended use by adding additional acetone or byallowing solvent to evaporate. For example, the mixture can be cast intorough form and machined into final form or cast into a preselected finalform and allowed to cure. Alternatively, suitable submicron powders canbe added to the mixture followed by additional blending. The paste canthen be applied as previously discussed.

The present invention has been described for use in the airfoil sectionof a narrow chord turbine blade. However, the present invention can finduse in other hot section components, such as liners, vanes, centerbodies and the like, as well as other sections of the blade such asplatforms and dovetails, in which small multiple plies are cut to sizeto account for a contour change or a thickness change, particularly overa short distance, and the substantially isotropic properties of adiscontinuously reinforced ceramic insert are adequate for theapplication. These applications are illustrated in FIG. 7. Two of theapplications are outside corners. FIG. 7A depicts the use of a ceramicinsert 710 of the present invention for use in a blade platform fillet,the insert 710 overlying the full length plies and replacing smallcorner plies. FIG. 7B depicts a similar use of the ceramic insert 730 asa replacement for small multiple ceramic plies along sharp outsidecorners, the insert 730 overlying the full length plies and replacingsmall corner plies. FIG. 7C depicts the use of a ceramic insert 750 ofthe present invention as a replacement for small multiple ceramic pliesfor a stiffener, in which there is a large change of contour alongcomponent cross-section, insert 750 replacing small cut plies andsurrounded by full length plies.

While the invention has been described with reference to a preferredembodiment, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment disclosed as the best modecontemplated for carrying out this invention, but that the inventionwill include all embodiments falling within the scope of the appendedclaims.

1. A composite component for use in high temperature applications,comprising: a lay-up of substantially continuous plies, each ply in thelay-up of substantially continuous plies having a plurality of towsextending substantially parallel to each other in an uncured matrixmaterial, each ply being positioned to that the tows extend at apreselected angle to the tows in an adjacent ply; a non-plydiscontinuously reinforced composite insert having substantiallyisotropic properties, the insert having a predetermined shape comprisingfiber-including material in a matrix material; wherein the insert ispositioned adjacent to the lay-up of plies at regions of contour changesand thickness changes of the component, the insert being in contact withat least one ply.
 2. The composite component of claim 1 wherein thethickness changes of the component include significant thickness changesover small distances, the small distances measured in a directionsubstantially transverse to the direction of changes in thickness. 3.The composite component of claim 1 wherein the insert further includesat least fiber-including material selected from the group consisting offiber, tow and chopped pre-preg plies.
 4. The composite component ofclaim 1 wherein the non-ply discontinuously reinforced material is fullycured.
 5. The composite component of claim 1 wherein the non-plydiscontinuously reinforced material is at least partially cured.
 6. Thecomposite component of claim 1 wherein the non-ply discontinuouslyreinforced material
 7. The composite component of claim 1 wherein thecontinuous plies are ceramic plies is an uncured material.
 8. Thecomposite material of claim 7 wherein the uncured material is a slurry.9. The composite material of claim 7 wherein the uncured materials is apaste.
 10. The composite material of claim 7 wherein the paste forms athixotropic composition further comprising polymer additions.
 11. Thecomposite material of claim 7 wherein the paste forms a thixotropiccomposition further comprising sub-micron powder.
 12. The compositecomponent of claim 1 wherein the non-ply discontinuously-reinforcedmaterial is a discontinuously reinforced ceramic matrix compositematerial.
 13. A hot section gas turbine engine component, comprising: acured, reinforced, ceramic matrix composite portion comprising aplurality of continuous tows extending substantially parallel to eachother in a matrix; a discontinuously reinforced composite portion havingsubstantially isotropic properties located at regions of contour changesand thickness changes of the component, the discontinuously reinforcedcomposite portion comprising discontinuous fiber-including material in amatrix material; wherein the discontinuously-reinforced compositeportion is adjacent to the reinforced ceramic matrix composite portioncomprising the plurality of continuous tows, and the discontinuouslyreinforced composite portion cured to the reinforced ceramic matrixcomposite portion.
 14. The hot section gas turbine component of claim13, wherein the cured, reinforced ceramic matrix composite portionfurther includes a plurality of groups of continuous tows in a matrix,the tows in each group extending substantially parallel to each other ina matrix and each group oriented at a preselected angle to the tows inat least one other group.
 15. The hot section gas turbine component ofclaim 13 wherein a first group of tows is oriented at a firstpreselected angle selected from the group consisting of 0°, +45°, −45°,90°, −45°, +45°, wherein the tows oriented at 0° are oriented parallelto an axis of the component, and the tows in the at least one othergroup adjacent to the first group are oriented at a second preselectedangle different than the preselected angle of the adjacent group, thesecond preselected angle selected from the group consisting of 0°, +45°,−45°, 90°, −45°, +45°.
 16. The hot section gas component of claim 13wherein the discontinuously-reinforced composite portion havingsubstantially isotropic properties located at regions of contour changesand thickness changes is located along a corner of a turbine component.17. The hot section gas component of claim 13 wherein thediscontinuously-reinforced composite portion having substantiallyisotropic properties located at regions of contour changes and thicknesschanges is positioned as a stiffener at a thickness change.
 18. The hotsection gas component of claim 13 wherein the discontinuously reinforcedcomposite portion comprising discontinuous fiber further includesfiber-including material selected from the group consisting of fiber,tow and chopped pre-preg plies.
 19. The hot section gas component ofclaim 18 wherein the fiber-including material has a size of from about0.1-1 inch.
 20. The hot section component of claim 13 wherein thediscontinuously-reinforced composite portion comprising discontinuousfiber-including material in a matrix material further includes a loadingof about 10% by volume to about 50% by volume of fiber-includingmaterial in matrix material.
 21. The hot section component of claim 13wherein the turbine component is a narrow chord turbine blade and thediscontinuously-reinforced composite portion having substantiallyisotropic properties is located in a trailing edge portion of the blade,a cured, reinforced, ceramic matrix composite portion bonded to thediscontinuously reinforced composite portion.
 22. The hot sectioncomponent of claim 13 wherein the turbine component is a narrow chordturbine blade and the discontinuously-reinforced composite portionhaving substantially isotropic properties and wherein the portion ispositioned at a change in contour adjacent an air passageway as a rib.